This invention relates to a jet propulsion exhaust nozzle. More particularly, the invention relates to a two dimensional wedge/translating shroud combination for providing a variable geometry nozzle exit for improved jet engine performance over a wide range of throttle power settings.
A maneuvering fighter airplane that must operate at subsonic and supersonic speeds requires a propulsion exhaust nozzle with a variable geometry to achieve high performance over a wide range of throttle power settings. Because of the high internal performance attainable with variable-geometry axisymmetric nozzles, this type of nozzle has generally been selected for past and present day jet-propelled aircraft designs. Another reason the variable geometry axisymmetric nozzle has been selected has been its low weight for a nozzle-engine item supplied by the engine manufacturers to be installed as a separate item in the airframe supplied by an airframe manufacturer.
The "round" axisymmetric nozzle used in some prior art engines exhibit a small throat/exit area at a typical cruise power setting which is similar to a convergent nozzle and which represents only about 20-25% of the maximum cross sectional area of the nozzle. The design of the external nozzle surface to close the remaining 80-75% of the nozzle area (called closure area) determines the drag level of the nozzle. An external nozzle surface with a steep boattail angle has significantly higher drag, especially as the speed of sound is approached, than a shallow boattail angle nozzle surface. Because of higher weight and higher rate of leak between the nozzle leaves where axisymmetric nozzles are segmented around the circumference, the shallow boattail nozzle is rejected in a trade off of performance for weight saving. It would therefor appear that the solution would be to make the maximum nozzle diameter smaller, hence a smaller engine, which leads to another reason for the large amount of closure area. However, supersonic flight speeds of current fighter aircraft vary up to about a Mach number of 2.5. To achieve these speeds the nozzle must be in the afterburner power setting, i.e., the nozzle geometry must be varied to a convergent-divergent nozzle with a throat area about twice the cruise throat area. A large expansion area ratio (A.sub.e /A.sub.t .apprxeq. 1.6) is necessary to achieve high performance at M = 2.5 supersonic speeds. Since the throat area in after-burner power setting is twice the cruise throat area and additional area is required for expansion area, the maximum area of the nozzle must be large compared to the throat area at the cruise power setting which unfortunately is generally where the aircraft operates.
At transonic speeds the acceleration of the flow over the steep nozzle boattail and, in multi-engine aircraft, in the channel between the engines/nozzles creates low static pressures on the nozzle-afterbody surfaces which enclose the closure area of the afterbody. In addition, some recent aircraft have twin vertical tails which can also cause adverse pressure gradients on the nozzle-afterbody in addition to the low static pressures on the steep nozzle bottail. Efforts to reduce the steep nozzle boattail, or closure area, by increasing the expansion area while holding the cruise throat area constant causes the peak nozzle performance to shift to a higher nozzle pressure ratio which would be higher than the engine operating pressure ratio. The end result of changing expansion area is lower internal performance in the nozzle off-design case for engine pressure ratios.
The traditional method to maximize thrust-minus-drag at transonic speeds is to minimize the resulting afterbody drag from the low static pressures over the closure area of an airplane model with wind tunnel tests using static tests to determine the internal nozzle performance with variation of expansion area ratio. The closure area of an airplane configuration may have horizontal-tail booms which are used to provide the most effective horizontal tail aft-mounts and/or used for supersonic area ruling. Unfortunately, the horizontal tail booms also have adverse effect at subsonic to transonic speeds. The closure area also includes interfairings which are used to fill the base between the nozzles. These interfairings must be tailored depending on the aircraft configuration and each is independent of the other.
It is therefore an object of the present invention to provide a realistic variable-geometry aircraft engine nozzle for use at high subsonic and supersonic flight speeds. A further object of the present invention is to provide an improved jet engine exhaust nozzle that obviates the problem of integrating the steep boattailed "round" geometry of the cruise nozzles into the aft-end of the aircraft fuselage. An additional object of the present invention is an improved jet engine nozzle serving to reduce the total afterbody drag at transonic speeds for multi-engine aircraft.